The present invention relates to an aircraft structure made of composite material, in particular for panels of fuselages of aeronautical structures, or for similar aircraft structures.
It is common knowledge that the aeronautical industry requires structures that on the one hand withstand the loads to which they are subjected, meeting high requirements on strength and stiffness, and on the other hand are as light as possible. A consequence of this requirement is the more and more extensive use of composite materials in primary structures, because with appropriate application of said composite materials it is therefore possible to achieve an important weight saving relative to a design in metallic material.
Integrated structures in particular have proved to be very efficient in this respect. We speak of an integrated structure when the different structural elements are manufactured in one step. This is another advantage of the use of composite materials because with their condition of independent layers that can be laid up in the desired form, they offer the possibility of increasing integration of the structure, which moreover often produces a cost saving—equally essential when competing in the marketplace—as there are fewer individual parts to be assembled.
Aircraft fuselages can be made up of several panels, these panels being composed of skin, stringers and frames. The skin is reinforced longitudinally with stringers to reduce its thickness and so as to be competitive in weight, whereas the frames prevent general instability of the fuselage and may be subject to local loading. Within an aircraft fuselage we may encounter other structural elements, such as beams, which serve as framing for open sections of the fuselage or serve for supporting the loads introduced by the floor of the cabin of said aircraft.
For the time being, panels of aircraft fuselages are manufactured separately. From one side, the skin reinforced with stringers is manufactured as a single piece by using ATL and FP techniques and a curing cycle in autoclave. From the other side, CFRP contour frames are manufactured separately in RTM (Resin Transfer Molding) by using two main preforms (C-shaped+L stiffener). One of the functions of this L stiffener is to reinforce the mousehole area of the frame (i.e., the recesses for the passage of the stringers), which is one of the weakest areas of the frame.
Once the frames are manufactured, they are assembled to the skin reinforced with stringers by means of mechanical joints, such as rivets. This involves manufacturing and assembly processes which are relatively costly.
In recent years much effort has been devoted to achieving an ever increasing level of integration in the production of aircraft structures in composite material, so as to avoid the aforementioned disadvantages.
US 2012/0034416 A1, referring to a frame and a method for producing such a frame, discloses a frame for reinforcing the hull of a craft, particularly an aircraft, comprising at least one frame element shaped in accordance with the curvature of the hull and having recesses at the side facing the hull for the passage of longitudinal beams of the craft, said frame element further having a flange facing the hull, said flange disposed in a region of at least one of said recesses and extending along an outline of said one recess, each frame element being integrally formed from a fiber-reinforced plastic composite material and the recesses being designed as integral cut-outs of each frame element.
U.S. Pat. No. 5,242,523 A discloses a method for forming and curing an intricate structure of criss-crossing composite stringers and frames that are bonded to a skin panel. A structure constructed in accordance with this invention would be well-suited for use as a portion of an aircraft fuselage, a boat hull, or the like. The method is preferably practiced by applying uncured composite stringers to an uncured composite sheet panel. This is followed by placing cured frames crosswise over the stringers. The frames have openings at the locations where they intersect with the stringers which enables the frames to come into direct contact with the skin along most of their length. During the forming and curing process, the stringers are covered with a plurality of cauls, and the entire assembly of skin panel, stringers, frames and cauls is subjected to a vacuum bagging and curing process. The cauls serve to maintain both part shape and to control the flow of resin within the stringers as they are cured. Further, they probably eliminate the need for intermediate protective materials between the vacuum bag and the stringers.
The angle currently used between the webs of the omega-shaped stringers and the skin (about 50°) makes it difficult to manufacture frames with a continuous flange.
Besides that, there are certain limitations regarding the angle currently used for the omega-shaped stringers. This angle cannot be lower than a value, in order to be sure that the distance x between the feet of an omega-shaped stringer is lower than the distance y between feet of different omega-shaped stringers (stringers pitch). Additionally, the distance x between the feet of an omega-shaped stringer cannot be so high in order to avoid manufacturing issues during the lay-up of the skin.
The present invention aims to solve the disadvantages that have arisen previously.